Patent Application
20220356555
The invention relates to an aircraft part, such as a turbine blade or distributor blade.
In a turbojet, the exhaust gases generated by the combustion chamber can reach high temperatures, for example greater than 1200° C., if not 1600° C. The turbojet parts, in contact with these exhaust gases, such as turbine blades for example, must thus be able to retain their mechanical properties at these high temperatures.
To this end, it is known practice to manufacture certain “superalloy” turbine parts. Superalloys are a family of high-resistance metal alloys that can function at temperatures relatively close to their melting points (typically 0.7 to 0.8 times their melting temperature).
Nevertheless, a superalloy part always has a limit operating temperature above which creep of the part is too great for the part to be used.
To this end, it is known practice to manufacture aircraft parts comprising one or more cooling channels. A fluid, such as a gas leaving a low-pressure compressor, can be introduced into the cooling channel(s). Its circulation thus allows the part to be cooled.
Nevertheless, the walls of the cooling channel(s) are sensitive to the environment. In particular these walls can be oxidised and/or corroded during use of the part, which decreases its usage time.
One aim of the invention is to is to propose a solution for manufacturing a turbine part comprising a cooling channel less sensitive to oxidation and/or corrosion than the cooling channels of the prior art.
This aim is achieved within the scope of the present invention as a result of a part comprising a substrate made of a nickel—based superalloy, the substrate having a first average mass fraction of one or more first elements chosen from hafnium, silicon and chromium, the substrate comprising at least one open cavity in the part and preferably a cooling channel, the substrate further comprising a surface layer at least partially forming the cavity, the surface layer having a second average mass fraction of the first element(s) which is strictly greater than the first average mass fraction.
The invention is advantageously completed by the following characteristics, taken individually or in any one of their technically possible combinations:
Another aspect of the invention is an aircraft turbine comprising a part according to the invention.
Another aspect of the invention is an aircraft comprising a part according to the invention.
Another aspect of the invention is a method for manufacturing an aircraft part according to the invention, comprising at least the following steps:
The invention is advantageously completed by the following characteristics, taken individually or in any one of their technically possible combinations:
Another aspect of the invention is a method for cooling an aircraft part, wherein the part complies with the invention, the method comprising a step of injecting a cooling fluid in the cavity.
Other characteristics, aims and advantages of the invention will emerge from the description which follows, which is purely illustrative and not limiting, and which should be read with reference to the appended drawings in which:
In the set of figures, similar elements have identical references.
The term “superalloy” refers to an alloy having, at high temperatures and high pressure, very good resistance to oxidation, corrosion, creep and cyclic constraints (especially mechanical or thermal). Superalloys have a particular application in the manufacture of parts used in aeronautics, for example turbine blade, as they constitute a family of high-resistance metal alloys which can work at temperatures relatively close to their melting points (typically 0.7 to 0.8 times their melting temperature).
A superalloy can have a biphasic microstructure comprising a first phase (called “γ phase”) forming a matrix, and a second phase (called “γ′ phase”) forming precipitates which harden in the matrix. The coexistence of these two phase is designated by the phrase “γ-γ′ phase”.
The “base” of the superalloy refers to the main metal component of the matrix. In most cases, superalloys comprise an iron, cobalt or nickel base, but also sometimes a titanium or aluminium base. The superalloy base if preferably a nickel base.
The “nickel-based superalloys” have the advantage of offering a good compromise between resistance to oxidation, resistance to rupture at high temperature and weight, which justifies their use in the hottest part of the turbojet.
Nickel-based superalloys consist of a γ phase (or matrix) of the γ—Ni face-centred cubic austenitic type, possibly containing additives in solid solution with a substitution (Co, Cr, W, Mo) and a γ′ phase (or precipitates) of the type γ′—Ni3X, with X=Al, Ti or Ta. The γ′ phase has an ordered L12 structure, derived from the face-centred cubic structure, coherent with the matrix, that is to say, having an atomic mesh very similar to it.
By virtue of its well-ordered nature, the γ′ phase has the remarkable property of having mechanical resistance which increases with the temperature up to about 800° C. The very strong coherence between the γ and γ′ phases confers a very high hot mechanical resistance to nickel—based superalloys, which is itself dependant on the γ/γ′ ratio and the size of the hardening precipitates.
A superalloy is preferably rich in rhenium and/or ruthenium, that is to say, the average mass fraction in rhenium and ruthenium of the superalloy is greater than or equal to 3%, and preferably to 4%, allowing increased creep resistance of the superalloy parts compared to superalloys without rhenium.
A superalloy is preferably generally poor in chromium, that is to say, the average mass fraction in chromium throughout the superalloy is less than 5%, preferably less than 3%. In fact, chromium depletion during rhenium and/or ruthenium enrichment of the superalloy makes it possible to retain a stable allotropic structure, especially a γ/γ′ phase,
The term “mass fraction” refers to the ratio of the mass of an element or group of elements to the total mass.
By “protective coating” is meant a layer covering the substrate and allowing it to be chemically and/or mechanically protected. The protective coating preferably prevents corrosion and/or oxidation of the substrate. Preferably, the protective coating may be a bonding layer between the substrate and a heat protection layer.
By “open cavity” of a part is meant a cavity linked to the exterior of the part.
By “secondary vacuum” is meant a vacuum in which the atmosphere is set at a pressure between 10−7 millibars and 10−3 millibars excluded.
By “primary vacuum” is meant a vacuum in which the atmosphere is set at a pressure between 10—3 millibars and 1 millibar.
Substrate 2
With reference to
The substrate 2 preferably has a low first average mass fraction in chromium throughout the substrate, that is to say, less than 5%. Thus, the substrate presents mechanical properties of creep resistance at a high temperature superior to a substrate with a first average mass fraction in chromium greater than 5%. Table 1 gives examples of the composition of the substrate 2, as an average mass fraction of each element throughout the substrate 2.
With reference to
Method for manufacturing the part 1 and protecting the cavity 12
With reference to
With reference to
The thickness /1of the layer 14 deposited during step 102 can be between 10 nm and 10 μm. When the first element is hafnium, the thickness /1 of the layer 14 deposited is preferably between 50 nm and 500 nm. When the first element is silicon, the thickness /1 of the layer 14 deposited is preferably between 100 nm and 500 nm. When the first element is chromium, the thickness /1 of the layer 14 deposited is preferably between 0.5 micrometres and 3 micrometres.
Depositing of the layer(s) 14 on the cavity 12 can be carried out by vapour phase chemical deposits (CVD), such as PECVD, LPCVD, UHVCVD, APCVD, ALCVD, UHVCVD.
With reference to
With reference to
Nonetheless, it is possible that, during the heat treatment step 203, certain elements of the substrate 2 are incorporated into the layer 14. Thus, the coating C2 has a mass fraction of the first element(s) greater than 50%, and preferably greater than 90%. The thickness /2 of the surface layer C1 is greater than 50 nm, that is, at the length characteristic of diffusion of the first element(s). The thickness /2 can in particular be greater than 100 nm, and preferably between 100 nm and 100 μm. The coating C2 has a thickness /3 between 50 nm and 100 μm.
Preferably, the surface layer C1 has a second mass fraction of the first element adapted to form a protective coating by oxidation of the first element. When the first element is hafnium, the second mass fraction can preferably be between 0.4% and 4.5%. When the first element is silicon, the second mass fraction can be preferably between 4% and 10%. When the first element is chromium, the second mass fraction can preferably be between 0.2% and 5%.
The substrate 2 and the layer(s) 14 obtained during step 202 may be, for example, placed in a chamber to carry out the heat treatment step 203. During the heat treatment step 203, the chamber can be put under vacuum, or filled with one or more inert gases, such as argon and/or helium. Preferably, a secondary vacuum can be maintained inside the chamber. Preferably, a primary vacuum can be set inside the chamber, the primary vacuum being made up of at least one element chosen from argon, helium and dihydrogen. Thus, it is possible to avoid oxidation of the surface of the substrate 2 during the heat treatment step 203. Preferably, the heat treatment step 203 comprises a temperature rise sub-step in which the temperature in the chamber is set so as to increase at a rate of 5 to 100° C. per minute. Preferably, the heat treatment step is carried out for one to eight hours, in a chamber in which the temperature is set between 700° C. and 1300° C., and preferably between 900° C. and 1250° C. Above 700° C., and preferably above 900° C., the first element(s) diffuse(s) into the substrate 2. The temperature is set below 1300° C., and preferably below 1250° C., so as not to degrade the superalloy.
| Number | Date | Country | Kind |
|---|---|---|---|
| 1912379 | Nov 2019 | FR | national |
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/FR2020/052002 | 11/5/2020 | WO |
